This invention relates generally to gas turbine engines, and more specifically to turbine blades used with gas turbine engines.
At least some known gas turbine engines include a core engine having, in serial flow arrangement, a high pressure compressor which compresses airflow entering the engine, a combustor which burns a mixture of fuel and air, and a turbine which includes a plurality of rotor blades that extract rotational energy from airflow exiting the combustor the burned mixture. Because the turbine is subjected to high temperature airflow exiting the combustor, turbine components are cooled to reduce thermal stresses that may be induced by the high temperature airflow.
The rotating blades include hollow airfoils that are supplied with cooling air through cooling circuits. The airfoils include a cooling cavity bounded by sidewalls that define the cooling cavity. Cooling of engine components, such as components of the high pressure turbine, is necessary due to thermal stress limitations of materials used in construction of such components. Typically, cooling air is extracted air from an outlet of the compressor and the cooling air is used to cool, for example, turbine airfoils. The cooling air, after cooling the turbine airfoils, re-enters the gas path downstream of the combustor.
At least some known turbine airfoils include cooling circuits which channel cooling air flows for cooling the airfoil. More particularly, internal cavities within the airfoil define flow paths for directing the cooling air. Such cavities may define, for example, a serpentine shaped path having multiple passes. Cooling air is directed through a root portion of the airfoil into the serpentine shaped path. In at least some known airfoil designs, an abrupt transition extends between the root portion and the airfoil portion to increase the cross-sectional area of the cooling cavity to facilitate increasing the volume of cooling air entering the airfoil portion. Because thermal stresses may be induced into the internal cavities, walls defining the cavities may be coated with a environmental coating to facilitate preventing oxidation within the cooling cavity. Because of the geometry of the cooling passages, during coating process, the coating is also deposited within the root portion of the airfoil.
To facilitate withstanding internal thermal stresses, at least some known blades are coated with a layer of environmental coating that has a thickness approximately equal to 0.001 inches. Applying the environmental coating with such a thickness prevents oxidation of the cavity walls and facilitates the airfoil withstanding thermal and mechanical stresses that may be induced within the higher operating temperature areas of the blade. However, if the coating is applied at a greater thickness, the combination of the increased thickness of the environmental coating and the abrupt transition within the dovetail may cause premature cracking in the root portion of the airfoil as stresses are induced into the transition area of the dovetail. Over time, continued operation may lead a premature failure of the blade within the engine.